1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air cooled turbine airfoil with trailing edge cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes a turbine section with one or more stages of stator vanes and rotor blades that react with a hot gas flow from a combustor to produce mechanical work and, in the case of an industrial gas turbine engine, drive an electric generator. It is known in the art that the engine efficiency can be increased by passing a higher temperature gas flow into the turbine. However, the turbine inlet temperature is limited by the material properties of the first stage airfoils and the amount of cooling provided for these airfoils.
Turbine airfoils are cooled by passing bleed off air from the compressor and through an internal cooling air passage within the airfoil. The cooling air from the compressor used for airfoil cooling is discharged from the airfoil without producing any useful work. Thus, the engine efficiency is reduced because the work used to compress the air used for airfoil cooling is lost. Therefore, it is also desirable to make use of a minimal amount of compressed air from the compressor used for airfoil cooling.
An airfoil is exposed to different temperatures due to the shape and the flow pattern across the airfoil. The hot gas flow strikes the leading edge of the airfoil and then flows around to the pressure side and the suction side. The trailing edge of the airfoil is the thinnest portion of the airfoil and is also exposed to some of the highest temperatures. Because of this, it is difficult to design for a cooling circuit for the trailing edge region. In the prior art, the trailing edge region of an airfoil is cooled by passing cooling air through channels that include pin fins to increase the heat transfer rate. FIG. 1 shows a prior art turbine airfoil for a first stage rotor blade with a row of drilled cooling air holes formed along the trailing edge of the blade. FIG. 2 shows a cross section view from the top of the FIG. 1 blade. The FIG. 1 design uses a single pass axial flow cooling channel to supply cooling air for the trailing edge region of the airfoil. The remaining sections of the airfoil are cooled with a separate serpentine flow cooling circuit. However, the single pass axial flow cooling design is not the best method for utilizing cooling air and therefore results in a low convective cooling effectiveness for the airfoil.